apollo 4 mission report

8/8/2019 Apollo 4 Mission Report

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MissionPA-I

A-001

AS-101

AS-I02

A-002

AS-103

A-003

AS-lOb

PA-2

AS-105

A-O04

AS-201

AS-202

Apollo 4

APOLLO SPACECRAFT FLIGHT HISTORY

Spacecraft Description Launch dateBP-6 First pad abort Nov. 7, 1963

BP-12 Transonic abort May 13, 1964

BP-13

BP-15

BP-23

BP-16

BP-22

BP-26

BP-23A

Nominal launch andexit environmentNominal launch andexit environmentMaximum dynamicpressure abort

MicrometeoroidexperimentLow-altitude abort(planned high-altitude abort)Micrometeoroidexperiment andservice moduleRCS launchenvironmentSecond pad abort

May 28, 1964

Sept. 18, 1964

Dec. 8, 1964

Feb. 16, 1965

May 19, 1965

May 25, 1965

June 29, 1965

BP-gA

SC-002

SC-009

SC-011

SC-017LTA-10R

Mierometeoroidexperiment andservice moduleRCS launchenvironmentPower-on tumblingboundary abort

Supercircularentry with highheat rateSupercircularentry with highheat loadSupercircularentry at lunarreturn velocity

July 30, 1965

Jan. 20, 1966

Feb. 26, 1966

Aug. 25, 1966

Nov. 9, 1967

Launch siteWhite SandsMissile Range,N. Mex.White SandsMissile Range,N. Mex.Cape Kennedy,Fla.

Cape Kennedy,Fla.White SandsMissile Range,N. Mex.Cape Kennedy,Fla.White SandsMissile Range,N. Mex.

Cape Kennedy,Fla.

White SandsMissile Range,N. Mex.Cape Kennedy,Fla.

White SandsMissile Range,N. Mex.Cape Kennedy,Fla.

Cape Kennedy,Fla.

Cape Kennedy,Fla.

_...A


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MSC-PA-R-68-1

APOLLO 4 MISSION REPORT

BY AUTHORITY OF

DATE SEP 2 0 19ZI

Prepared by: Apollo 4 Mission Evaluation Team

Approved by:ManagerApollo Spacecraft Program

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIO_MANNED SPACECRAFT CENTER

HOUSTON, TEXASJanuary 7, 1968

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APOLLO 4 SPACE VEHICLE.


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iv

Sect ion5.2

5.3

5.2

5.5

5.6

AL

AERODYNAMICS .................5.2. i Summary ...............5.2.2 Aerodynamics Predictions .......5.2.3 Flight-Derived Aerodynamics .....5.2.2 Comparison of Predictions With

Flight-Derived Aerodynamic DataTHERMAL STRUCTURES ..............5.3.1 Launch Phase .............5.3.2 Orbital Flight ............AEROTHERMODYNAMICS AND HEAT PROTECTION

SUBSYSTEM .................5.2.1 Aerot hermodynamic s ..........5.2.2 Heat Protection Subsystem ......EARTH LANDING SUBSYSTEM (ELS) ........5.5.1 Performance .............5.5.2 Postflight Test Activity .......MECHANICAL SUBSYSTEMS ............5.6.1 Summary ...............5.6.2 Performanc e .............ELECTRICAL POWER SUBSYSTEM ..........FUEL CELLS ..................5.8.1

5.8.25.8.3

Summary ...............Prelaunch Operation .........Performance .............

C TIAL

Page5.2-15.2-15.2-15.2-2

5.2-25.3-15.3-15.3-3

5._-i5.4-15.2-65.5-15.5-15.5-25.6-15.6-15.6-15.7-15.8-15-8-1

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Section5.9 SEQUENTIAL EVENTS CONTROL SUBSYSTEM .....

5.10 PYROTECHNIC SUBSYSTEM ............5.11 LAUNCH ESCAPE SUBSYSTEM ...........5.12 EMERGENCY DETECTION SUBSYSTEM ........

5.12.1 Summary ...............5.12.2 Description .............5.12.3 Performance .............

5.13 COMMUNICATIONS SYSTEM PERFORMANCE ......5.13.1 Summary ...............5.13.2 Communications Subsystem

Performance ............5.13.3 Performance During Launch Phase5.13.h Performance During Near-Earth

Parking Orbit ...........5.13.5 Performance During Translunar

Injection and High Ellipse Phase5.1h INSTRUMENTATION ...............5.15 GUIDANCE AND CONTROL SUBSYSTEM ........

5.15.1 Summary ...............5.15.2 Integrated Subsystem Performance5.15.3 Guidance and Navigation Subsystem

Performance ............

5.15.h Stabilization and ControlSubsystem Performance .......5.15.5 Mission Control Programmer

Performance ............

TIA L

Pag e5.9-15.10-15. ll-15.12-I5.12-15.12-15.12-i5.13-15.13-1

5.13-8

5.13-115.1h-15.15-15.15-15.15-1

5.15-7

5.15-10

5.15-ii

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vi

Section

t

5.16 REACTIONONTROLUBSYSTEM.........5.16.1 Service Module Reaction Control

Subsystem .............5.16.2 CommandModule Reaction ControlSubsystem .............5.17 SERVICEPROPULSIONUBSYSTEM........

5.17.1 SubsystemDescription ........5.17.2 Propellant Loading ..........5.17.3 Service Propulsion SubsystemMissionDescription ............5.17.h Steady-State Performance andAnalysis .............5.17.5 Normalized Performance ........5-17.6 Gauging SubsystemAnalysis ......5.17.7 Pressurization Subsystems ......5.17.8 Engine Transient Analysis ......

5.18 CRYOGENICUBSYSTEM............5.18.1 Summary ...............5.18.2 Prelaunch Operations ........ .5.18.3 Performance .............

5.19 ENVIRONMENTALONTROLUBSYSTem4......5.19.1 Launch Phase ............5.19.2 Earth Orbital Phase .........5.19.3 Entry Phase .............5.19.h Postrecovery Observations ......

C

Page5.16-i

5.16-15.16-65.17-15.17-15.17-2

5.17-2

5.17-35.17-55.17-55.17-65.17-75.18-15.18-15.18-i5.18-25.19-i5.19-15.19-25.19-35.19-h


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vii

Section

6.0

7.08.09.0

5.20 CREW STATION PERFORMANCE ...........

5.20.1 Crew Visibility ...........5.20.2 Crew Related Dynamics ........5.20.3 Radiation Monitoring .........

LUNAR MODULE TEST ARTICLE .............6.1 STRUCTURE ..................

6. i. 1 Loads ................6.1.2 Low Frequency Vibration .......6. i. 3 Vibro-Acoust ics ...........

6.2 INSTRUMENTAT ION ...............FLIGHT CREW ....................BIOMEDICAL EVALUATION ...............MISSION SUPPORT ........... .......9.1 FLIGHT CONTROL ................

9. i. 1 Prelaunch ..............9. i. 2 Launch Phase .............9.1.3 Earth Orbit - Revolution 1 ......9.1.4 Earth Orbit - Revolution 2 ......9.1.5 Translunar Injection - Revolu-

tion 3 ..............9. i. 6 Entry ................

9.2 NETWORK PERFORMANCE .............9.2. i Apollo/Range Instrumentat ion

Aircraft .............

CON

Page5.20-15.20-15.20-35.2o-46-16-16-16-26-36-217-18-19-19-19-19-39-49-5

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Section

i0.0ii. 012.0

CONFPage

9.2.2 Telemetry .............. 9-119.2.3 Tracking ............... 9-119.2.4 Command ............... 9-129.2.5 Goddard Space Flight Center Central

Processors .............Real Time Computer Complex .......2.6

9-139-139-159-159-16

9.3 RECOVERY OPERATIONS .............9.3.1 Recovery Force Deployment ......9.3.2 Spacecraft Location and Retrieval9.3.3 Recovery Force Electronic

Reception ............. 9-189.3.4 Spacecraft Recovery Inspection .... 9-209.3.5 Spacecraft Deactivation ....... 9-229.3.6 S-IC and Camera Capsule Recovery . . 9-22

EXPERIMENTS .................... i0-iCONCLUSIONS ..................... ii-iANOMALY SUMMARY AND POSTFLIGHT TESTING ....... 12-i12.1 MISSION ANOMALIES .............. 12-112.2 COUNTDOWN ANOMALIES ....... _ ..... 12-612.3 TEST AND CHECKOUT ANOMALIES ......... 12-712.4 POSTFLIGHT TESTING .............. 12-8

12.4.1 Heat Protection Subsystem ...... 12-812.4.2 Earth Landing Subsystem ....... 12-8

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Section

13.0

C TIA L ixPage

12.h. 3 Mechanical Subsystems ......... 12-8

12.h. h Electrical Power Subsystem ...... 12-812.h. 5 Environmental Control Subsystem .... 12-912.4.6 Communications ............ 12-912.h.7 Pyrotechnics ............. 12-912.4.8 Instrumentation ............ 12-912.4.9 Guidance and Control Subsystems .... 12-1012.h.lO Reaction Control Subsystem ...... 12-10

VEHICLE AND SYSTEMS DESCRIPTION ........... 13-113.1 COMMAND AND SERVICE MODULE .......... 13-3

13.1.1 Structures .............. 13-313.1.2 Earth Landing System ......... 13-413.1.3 Mechanical Subsystem ......... 13-513.1.h Electrical Power Subsystem ...... 13-713.1.5 Sequential Events Control Subsystem . 13-913.1.6 Pyrotechnic Devices .......... 13-1013.1.7 Emergency Detection Subsystem ..... 13-1013.1.8 Conmunications Subsystem ....... 13-1113.1.9 Instrumentation Subsystem ....... 13-1113.I.I0 Guidance and Control ......... 13-1213.1.11 Reaction Control Subsystem ...... 13-1513.1.12 Service Propulsion Subsystem ..... 13-1713.1.13 Environmental Control Subsystem .... 13-18

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Section

1h.0

15.016.0

Page13.2 LUNAR MODULE TEST #_TICLE ........... 13-40

13.2.1 General Description .......... 13-4013.2.2. Instrumentation and Co_nunication . . . 13-40

13.3 LAUNCH VEHICLE DESCRIPTION .......... 13-4313.h WEIGHT AND BALANCE DATA ............ 13-47SPACECRAFT HISTORIES ................ lh-ilh.l COMMAND MODULE AND SERVICE MODULE ....... lh-i1h.2 LUNAR MODULE TEST ARTICLE ........... lh-h

REFERENCES ..................... 15-1DISTRIBUTION .................... 16-1

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Table2.1-I2.2-I3.1-I

3.1-II

3.l-III

3.1-IV3. l-V3. I-VI

3.1-VII3.1-VIII3.l-IX

3.1-X

5.1-I5.1-II5.1-III

5.1-IV

5.2-I

TABLES

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Page2-32-5

APOLLO 4 MISSION EVENTS ..............DETAILED TEST OBJECTIVES .............LAUNCH PHASE PLANNED AND ACTUAL TRAJECTORY

PARAMETERS ................... 3-7PARKING ORBIT PLANNED AND ACTUAL TRAJECTORY

PARAMETERS ................... 3-9ORBITAL ELEMENTS ................. 3-11

REVOLUTION 1 TRACKING RESIDUAL STATISTICS ..... 3-12REVOLUTION 2 TRACKING RESIDUAL STATISTICS ..... 3-13COAST ELLIPSE PLANNED AND ACTUAL TRAJECTORY

PARAMETERS ................... 3-14COAST ELLIPSE TRACKING RESIDUAL STATISTICS .... 3-18SPS THRUST AND TARGETING CHARACTERISTICS ..... 3-19ENTRY INTERFACE PLANNED AND ACTUAL TRAJECTORY

PARAMETERS ................... 3-20SUMMARY OF PLANNED AND ACTUAL TRAJECTORY

PARAMETERS .................MAXIMUM LATERAL LOAD CONDITIONS ..........MAXIMUM SPACECRAFT LOADS AT MAXIMUM qa ......MAXIMUM SPACECRAFT LOADS AT END OF FIRST-STAGE

BOOST ......................LAUNCH ESCAPE SYSTEM AND COMMAND MODULE LOW

FREQUENCY VIBRATION DURING FIRST-STAGELAUNCH AND BOOST ................

COMMAND MODULE AERODYNAMICS-PREFLIGHTPREDICTIONS ................... 5.2-4

3-215.i-i05.i-ii

5.1-12

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Table5.4-I

5.4-II

5.4-III5.4-IV5.8-I

5.13-1

5.15-III

5.15-IV

5.15-V

5.15-VI5.15-VII

5.15-VIII

5.15-IX

LOCATIONS AND RANGES OF PRESSURE SENSORS,CALORIMETERS, AND RADIOMETERS ..........

TABLE OF HEAT SHIELD COMPONENT AND EQUIPMENTTHERMAL RESPONSE ................AFT HEAT SHIELD ABLATOR DATA ...........HEAT SHIELD BONDLINE T_4PERATURES .........SUMMARY OF FUEL CELL HYDROGEN COMSUMPTION AND

WATER PRODUCTION ................COMMAND EVENTS - BERMUDA AND THE

U.S.N.S. VANGUARD ................STATE VECTOR COMPARISON ..............SERVICE PROPULSION SUBSYSTem4 GIMBAL TRIM

VALUES ........... ..........COMPARISON OF ATTAINED ORBIT WITH TARGET DURING

FIRST SPS BURN .................COMPARISON OF ATTAINED ORBIT WITH TARGET DURING

SECOND SPS BURN .................APOLLO GUIDANCE COMPUTED ENTRY NAVIGATION AND

GUIDANCE RECONSTRUCTION .............PRELIMINARY IMU ERROR COEFFICIENTS ........APOLLO GUIDANCE COMPUTER ENTRY NAVIGATION

ACCURACY ....................INERTIAL MEASUREMENT UNIT PREFLIGHT PERFORMANCE

SUMMARY .....................COMPARISON OF INFLIGHT ACCELEROMETER BIAS WITH

PREFLIGHT COMPENSATION .............

APOLLO GUIDANCE COMPUTER MAJOR MODE TIMELINESCS PARAMETER COMPARISON .............

CON , IAL

Page

5.4-12

5.4-155.4-185.4-19

5.8-4

5.13-175.15-13

5.15-14

5.15-15

5.15-16

5.15-175.15-18

5.15-19

5.15-20

5.15-215.15-225.15-24

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Table5.16-I

5.16-II5.16-III

5.16-IV

5.16-VII

5.16-VIII5.16-IX5.16-X

5.16-XI

5.17-III

IALHELIUM AND PROPELLANT SERVICING ..........

RCS SEQUENCE OF EVENTS ..............TYPICAL MANEUVER ACCELERATIONS AND TRANSLATION

VELOCITY CHANGES ................SERVICE MODULE REACTION CONTROL SUBSYSTEM

PROPELLANT CONSUMPTION, ENGINE BURN TIMES,NUMBER OF PULSES, DUTY CYCLES, AND PRESSUREDROPS PER MANEUVER ...............

SM-RCS LAUNCH HEATING SUMMARY ...........SUMMARY OF SM-RCS TIMPERATURES FROM CSM/S-IVB

SEPARATION THROUGH CM/SM SEPARATION .......THERMAL CONTROL SUBSYSTf_4 PERFORMANCE SUMMARY

DURING COLD-SOAK PERIOD (03:29:00 TO08:01:00) ....................

TYPICAL CM-RCS CONTROL CROSS-COUPLING EFFECTS .CM-RCS PRESSURE TRENDS ..............COMMAND MODULE REACTION CONTROL SUBSYSTEM

PROPELLANT CONSUMPTION, ENGINE BURN TIMES,NUMBER OF PULSES, DUTY CYCLES, AND PRESSUREDROPS PER MANEUVER ...............

COMMAND MODULE REACTION SUBSYSTEM THERMALPERFORMANCE SUMMARY ...............

FLIGHT DATA USED IN THE ANALYSIS PROGRAM .....SERVICE PROPULSION SUBSYSTEM PERFORMANCE SUMMARY,

APOLLO 4 MISSION ................SERVICE PROPULSION SUBSYSTEM ENGINE TRANSIENT

ANALYSIS SUMMARY ................SPACECRAFT WINDOW RESOLUTION CHARACTERISTICS . .LUNAR MODULE LOW FREQUENCY VIBRATION DURING FIRST

STAGE LAUNCH AND BOOST .............L

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5.16-13

5.16-145.16-18

5.16-19

5.16-205.16-215.16-22

5.16-23

5.17-9

6-7


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Table9.3-I12.4-I13.4-I

13.4-II

RECOVERY SUPPORT .................ASHUR's FOR POSTFLIGHT ANOMALY TESTING ......SPACECRAFT MASS PROPERTIES AT LAUNCH AND DURING

ORBITAL FLIGHT .................COMMAND MODULE MASS PROPERTIES AT ENTRY ......

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TIALFIGURES

XV

Figure3.1-13.1-2

3.1-3

3.1-4

3.1-5

5.1-1

Apollo h mission ground track ...........Time history of the launch phase(a) Latitude, longitude, and altitude ......(b) Space-fixed flight-path angle

and velocity ................(c) Earth-fixed flight-path angle(d) Mach number and dynamic pressure .......Orbital time history of space-fixed flight-path

angle, velocity, and altitude(a) Parking orbit ................(b) Coast ellipse ................Service propulsion subsystem burn time history of

space-fixed flight-path angle, velocity, andaltitude(a) First service propulsion subsystem burn . . .(b) Second service propulsion subsystem burn .

Time history of entry phase(a) Latitude, longitude, altitude ........(b) Space-fixed flight-path angle and

velocity ..................(c) Earth-fixed flight-path angle and

velocity ..................(d) Load factor .................Launch escape system and command moduleaccelerometer locations .............S-IC engine thrust build-up ............Spacecraft accelerations at lift-off(a) LA0011A, LA0012A, and CAO001A ........(b) CA0007A, CK0036A, and CK0037A ........

C0 1A L

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3-233-24O r)_j--m.j3-26

3-273-28

3-293-3O

3-313-323-333-34

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Figure5.1-4

5.1-5

5.1-6

5.1-7

5.1-8

5.1-9

5.1-10

5. i-ii

5.1-12

5.1-13

CO TIAL

Launch winds

(a) Launch winds magnitude ............(b) Launch winds direction ............Comparison of maximum qa command module/service

module interface loads with structuralcapabilities ..................

Comparison of maximum qa service module/spacecraftlunar module adapter interface loads withstructural capabilities .............

Comparison of maximum qa spacecraft lunarmodule/instrument unit interface loads withstructural capabilities .............

Launch escape Zower lateral oscillations duringfirst-stage boost ................

Spacecraft accelerations during high axialacceleration oscillation period

(a) LA0011A, LA0012A, and CA0001A ........(b) CA0007A, CA0036A, and CK0037A ........Spacecraft accelerations at first-stage end of

boost(a) LA0011A, LAO012A, and LA0001A ........(b) CA0007A, CK0036A, and CK0037A ........Spacecraft accelerations during S-IC/S-II separation(a) LA0011A, LA0012A, and CA0001A ........(b) CA0007A, CK0037A, and CK0036A ........Typical command module/service module tension tie

strain gauge location ..............

SLA structural measurement locations .......

Page

5.1-20

5.1-21

5.1-22

5.1-23

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Figure5.1-14

5.1-15

5.1-16

5.1-17

5.2-3

5.2-h5.3-1

5.3-2

5.3-3

5.3-4

5.3-5

Command module acceleration spectral density(a) Prior to ignition ...............(b) Maximum vibration from 85.0 to 87.5 seconds .Comparison of command module internal equipment

bay's (Block I) vibration criteria to measuredvibration during atmospheric flight .......

Service module acceleration spectral density(a) Prior to ignition ...............(b) Maximum vibration from 77.0 to 79.0 secondsComparison of service module aft bulkhead vibration

criteria to measured vibration during atmosphericflight ......................

Command module external configuration .......Aerodynamic correlation parameters(a) Mach number ..................(b) Reynolds number per foot ...........Command module total lift-to-drag ratio ......

Command module total angle-of-attack ........Location of service module inner skin

thermocouples ..................Peak temperatures measured on spacecraft lunar module

adapter outer surface during the launch phase .Temperature measured by spacecraft lunar module

adaptor sensor AA7863T ..............Temperature measured by spacecraft lunar module

adaptor sensor AA786hT ....... .......Outer surface temperatures of the spacecraft lunar

module adapter at S-IC outboard engine cutoff

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Page

5.1-34

5.2-8

5.2-9

5.3-6

5.3-7

5.3-8

5.3-9

5.3-10

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xviii C TI"ALFigure5.3-6

5.3-75.3-8

5.3-9

5.3-10

5.4-1

5.4-4

5.4-5

5.4-6

Temperature sensor locations on command moduleand predicted maximum temperatures ........

Forward and side heat shield measurement locationsand values just prior to entry ..........Spacecraft attitude referenced to the sun during

cold-soak period .................Temperature measured by command module sensor

CAI502T during cold-soak period .........Temperature measured by command module sensor

CAI509T during cold-soak period .........Sketch of Apollo command module showing locations

of aerothermodynamic sensors ...........Free-stream density ................Pressures measured on aft heat shield(a) Location i ..................(b) Location 2 . . . . .............(c) Location 3 .................(d) Location 4 ..................(e) Location 5 .................(f) Location 6 ..................(g) Location 7 ..................(h) Locations 12, 13, and 16 ...........(i) Locations 18, 19, and 20 ...........Distribution of measured pressure ratio with wind

tunnel data a = 25 ...............Comparison of measured radiative heating rate with

theoretical predictions(a) First heat pulse ....... ........(b) Second heat pulse ..............

Total radiative heating rate prediction with itscomponents ...................

c TIAL

Page

5.3-11

5.3-12

5.3-13

5.3-14

5.3-15

5.4-225.4-235.4-245.4-255.4-265.4-275.4-285.4-295.4-30

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C' ENTJAL xixFigure5.4-7

5.4-8

5.4-9

5._-zo

5.4-lZ

5.4-12

PageHigh range calorimeter wafer temperatures(a) Locations I, 2, 5, and 6 ...........(b) Locations 8, 9, and ll ............(c) Top 4 wafers of location 3 ..........(d) Top 3 wafers of location 7 ..........(e) Top 3 wafers of location 8 ..........

5.4-355.4-365.4-375.4-385.4-39

Comparison of aft heat shield heating rate obtained fromwafer temperature measurement with theoretical prediction

(a) Location i ..................(b) Location 2 ..................(c) Location 3 ..................(d) Location 5 ..................(e) Location 6 ..................(f) Location 7 ..................(g) Location 8 ..................(h) Location l0 ..................

I I A5 4-4U5.4-hl5.4-425.4-435.4-445.4-455.4-465.4-47

Comparison of measured heating rate on the conicalsection with theoretical predictions

(a) Locations 12, 13, and 14 ...........(b) Locations 16, 17, 21, and 22 .........(c) Location 18 ..................(d) Locations 19 and 20 ..............(e) Locations 24, 25, 26, and 27 .........(f) Locations 15, 23, and 28 ...........(g) Singularity locations .............

5.4-485.4-495.4-505.4-515.4-525.4-535.4-54

Aft heat shield ablator temperature and charmeasurements .................. 5.4-55

Conical heat shield ablator and astro-sextant areatemperature and char measurements ........ 5.4-56

Major component thermodynamic measurement locations(a) Umbilical ...................(b) Parachute compartment .............(c) Launch escape system tower well anddisconnect .................(d) S-band window .................(e) Crew compartment (hatch stringer 12-C) ....

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XX

Figure

5.4-13

5.4-14

5.4-15

5.4-16

5.4-17

5.4-18

5.4-19

5.4-20

TIALPage

(f) Aft heat shield (tension tie i) ........(g) Command module air vent ............(h) Steam vent ..................(i) Command module stringer ............(j) Sidewall heat shield attachment ........(k) Command module window .............(1) C-band antenna ...... . .........

5.4-605.4-615.4-615.4-625.4-625.4-635.4-63

Postflight photograph showing char condition ofaft heat shield ................. 5.4-64Aft heat shield temperature measurements at depthsindicated(a) Station Z = +71.82, Y = 0 ...........(b) Station Z = -50.0, Y = -1.99 .........(c) Station Z = -64.85, Y = -14.68 ......

5.4-655.4-665.4-67

Aft heat shield 1050 F isotherm comparison with charsensor and core char ............. 5.4-68

Maximum temperature measured in depth and comparison ofchar interface with 1050 F isotherm

(a) Y = 0, Z = +71.82 ..............(b) Y = 0, Z= +5O ................Postflight photograph showing details of typical shear

compression pad, located at = 152 45' .... 5.4-70Postflight photograph showing char condition of

umbilical ramp .................. 5.4-71Toroidal heat shield temperature measurements at

depths indicated(a) Station X = 18.5, @ = 180 .......... 5.4-72C(b) Station X = 18.5, @ = 225 .......... 5.4-73C(c) Station X = 18.5, @ = 270 .......... 5.4-74CPostflight photographs of conical heat shield(a) +Z-axis to +Y-axis ..............(b) +Y-axis to -Z-axis ..............

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Figure

5.4-21

5.4-22

5.4-23

5.4-265.5-1

5.5-2

C DENTIAL

(c) -Y-axis to +Z-axis ..............(d) -Z lee side of spacecraft ...........(e) Forward compartment heat shield ........Conical heat shield temperature measurements at depths

indicatedCa)(b)Co)(d)(e)(f)(g)(h)(i)(j)

Station X = 26, = 90 ...........CStation X = 50, = 90 ...........cStation X = 84, = 90 ...........CStation X = 104, = 90 ...........CStation X CStation XcStation X cStation X cStation X cStation X c

= 26, = 135 ...........= 50, = 180 ...........= 80, = 180 ...........= I04, = 180 ..........= 80, = 270 ...........= 104, = 270 ..........

Comparison of Apollo 4 measured and predicted heatshield temperature data at X = 50 , = 180 usingc0.06-inch thermocouple driver ..........

Simulated windward umbilical cavity, located at = 87 .....................Charred astro-sextant area .............Unified side hatch test panel showing gap and

seal .......................Steam vent, air vent, and EVA hand hold ......Pressure altitude during descent ..........

Earth landing subsystem upper deck after recovery(a) Drogue mortar can no. i ............(b) Drogue mortar can no. 2 ...........

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5.4-805.4-815.4-825.4-835.4-845.4-855.4-865.4-875.4-885.4-89

5.4-9o

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xxi i C N ALFigure5.7-1

5.7-2

5.7-3

512-2

5.12-3

5.12-4

5.13-3

5.13-4

5.13-5

5.13-6

Main DC Bus A voltage and total spacecraft currentduring gimbal motors start-up ..........

Entry battery currents during gimbal motorsstart-up .....................

Entry battery voltages during gimbal motorsstart-up .....................

Load profile ....................Nominal fuel cell performance compared with

actual flight data ................Fuel cell no. 3 temperatures ............Emergency detection subsystem, angle of attack

parameter (AP) ..................Angular rates and emergency detection system abort

limits during S-II stage burn ..........Angular rates and emergency detection system abort

limits during first S-IVB stage burn .......Angular rates and emergency detection system abort

limits during second S-IVB stage burn ......Apollo 4 communications capabilities ........Uplink received carrier power at spacecraft, launch

phase ......................Downlink received carrier power at Manned Spaceflight

Network sites, launch phase ...........Downlink received carrier power at MIL, launch

phase ......................Downlink received carrier power at GBM, launch

phase ......................Downlink received carrier power at BDA, launch

phase ......................

C( I NTIA L.j

Page

5.7-2

5.7-3

5.12-3

5.12-4

5.12-5

5 13-20

5.13-21

5 13-22

5.13-23

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CO F TIALFigure5.13-7 Uplink received carrier power at spacecraft,

Hawaii coverage of second revolution .......5.13-8 Downlink received carrier power at Hawaii,

second revolution ................5.13-9 Downlink received carrier power at MIL, first

and second revolution ..............5.13-10 Unified S-band telemetry bit error probability,

Hawaii coverage of second revolution .......5.13-11 VHF telemetry bit error probability, Hawaii

coverage of second revolution ..........

5.13-12 Uplink received carrier power at spacecraft,ACN and CRO coverage of third revolution .....5.13-13 Downlink received carrier power at ACN and CRO,

third revolution ................5.13-14 Uplink received carrier power at spacecraft,

Guam coverage of third revolution ........5.13-15 Downlink received carrier power at Guam,

third revolution .................5.13-16 Uplink received carrier power at spacecraft,

A/RIA 3 coverage of entry ............5.13-17 Unified S-band telemetry bit error probability,

ascension coverage of third revolution ......5.13-18 Unified S-band telemetry bit error probability,

Guam coverage of third revolution ........5.15-1 Boost monitor phase, pitch error ..........5.15-2 Spacecraft dynamics, ascent phase .........5.15-3 S-IVB separation and service propulsion subsystem

burn number 1 sequence of events ........5.15-4 Velocity accumulation, S-IVB/CSM separation ....

CO ,., .

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Page

5.13-25

5.13-26

5.13-27

5.13-28

5.13-29

5.13-30

5.13-31

5.13-32

5.13-33

5.13-34

5.13-35

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xxi v

Figure5.15-5

5.15-6

5.15-7

5.15-8

5.15-9

5.15-10

5.15-11

5.15-12

5.15-13

5.15-! 4

5.15-15

5.15-16

5.15-17

Spacecraft dynamics, S-IVB/CSM separationand maneuver to service propulsion subsystemfirst burn attitude ...............

Spacecraft dynamics, service propulsionsubsystem first burn ...............

Velocity to be gained (VG) , service propulsionsubsystem first burn ...............

Spacecraft dynamics, maneuvers to cold-soakattitude .....................

Spacecraft dynamics, typical coast phase limitcycle ......................

Spacecraft dynamics, maneuver to service propulsionsubsystem second burn attitude ..........

Sequence of events for service propulsion subsystemsecond burn ...................

Spacecraft dynamics, service propulsion subsystemsecond burn ...................

Velocity to be gained (VG) , service propulsionsubsystem second burn ..............

Spacecraft Dynamics, entry(a) 08:18:00 to 08:22:00 .............(b) 08:22:00 to 08:26:30 . ............(c) 08:26:30 to 08:31:00 .............(d) 08:31:00 to 08:35:40 .............Entry sequence plotted against altitude and

range .......................Roll command plotted against actual roll (CDUX) .

Landing point data .................

Page

5.15-29

5.15-30

5.15-31

5.15-32

5.15-33

5.15-34

5.15-35

5.15-36

5.15-37

5.15-385.15-395.15-405.15-41

5.15-425.15-43

5.15-44

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FigureC IAL

5.15-18 Time histories of guidance and navigationsubsystem inertial instrument coefficients errors

(a) X, Y, and Z-axis accelerometer bias andscale factor error coefficients .......

(b) X, Y, and Z-axis gyro acceleration drift spinreference axis (ADSRA) and null bias driftcoefficients (NDB) .............

(c) X, Y, and Z-axis gyro acceleration drift inputaxis (ADIA) coefficient .... ......

5.15-19 Ascent velocity residuals .............5.16-i Average number of pulses accumulated during service

module reaction control subsystem activity ....5.16-2 Average burn time accumulated during service module

reaction control subsystem activity .......5.16-3 Service module reaction control subsystem propellant

consumption(a) Quad A consumption .............(b) Quad B consumption ..............(c) Quad C consumption ..............(d) Quad D consumption ..............(e) Total consumption ...............

5.16-4 Service module reaction control subsystem enginemounting structure temperatures(a) Quads A and B .................(b) Quads C and D .................

5.16-5

5.16-6

Service module reaction control subsystem engineinjector head temperatures ............

Quad B engine mounting structure temperature duringportion of the cold-soak period .........

5.16-7 Effect of engine firing on injector head and enginemounting structure temperatures .........

5.16-8 Command module reaction control subsystem heliumtank pressures during entry ...........

XXV

Page

5.15-45

5.15-465.15-475.15-48

5.16-25

5.16-26

5.16-275.16-285.16-295.16-305.16-31

5.16-325.16-33

5.16-34

5.16-35

5.16-36

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xxvi

Figure5.16-9

5.16-10

5.16-11

5.16-12

5.16-13

5.16-12

5.17-1

5.17-2

5.18-i5.19-15.19-25.20-1

5.20-2

Command module reaction control subsystem heliumtank temperatures during entry ..........

Typical command module reaction control subsystemengine chamber pressure for burns of 0.005,0.145 and 42.1 seconds ..............

Average number of pulses accumulated during commandmodule reaction control subsystem activity ....

Average burn time accumulated during command modulereaction control subsystem activity .......

Command module reaction control subsystem propellantconsumption(a) System A consumption .............(b) System B consumption .............(c) Total command module reaction control

subsystem ..................Typical command module reaction control subsystem

engine component temperatures during entry ....Service propulsion subsystem chamber pressure data(a) First burn ..................(b) Second burn ..................Service propulsion subsystem second burn

acceleration match ...............Oxygen tank pressure and events ..........Waste water tank quantity plotted against time . . .Potable water tank quantity plotted against time .Postflight photograph of left-side window after

spacecraft had been secured on recovery ship . . .Shipboard photographs showing postflight condition of

spacecraft 017 windows .......... . ....

Page

5.16-38

5.16-39

5.16-40

5.16-21

5.16-44

5.16-45

I5.17-115.17-12

5.17-135.18-45.19-65.19-7

5.20-7

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\

Figure5.20-3

5.20-4

5.2o-5

5.20-6

5.20-7

5.2O-8

5.20-9

6.1-5

6.1-6

6.1-7

6.1-8

xxvi i

PagePostflight photographs of spacecraft windows

taken after spacecraft arrived at Downey,California ....................

Postflight photographs of spacecraft windowsshowing discoloration due to contaminate .....

Postflight grid photographs of windows takenafter spacecraft arrived at Downey, California . .

Postflight resolution photographs taken afterspacecraft arrived at Downey, California .....

Acceleration input to command module X-axis atlift-off, measurement CAOOOIA ..........

Spectral density derived from X-axis accelerometerat time of lift-off, measurement CA00OIA .....

Spectral density derived from X-axis accelerometerduring period of maximum response, measurementCAOOOIA ..................... 5 20-15

+Z outrigger strut load typical 5 Hz oscillation . . 6-8Effect of 5 Hz oscillation on +Z apex X-component

of load ..................... 6-9

Lunar module accelerations at lift-off ....... 6-10Lunar module accelerations during high oscillation

period ...................... 6-11Lunar module accelerations during first stage

center engine cutoff ............... 6-12Lunar module accelerations during first stage

outboard engine cutoff .............. 6-13Acceleration spectral density of LTA-IOR oxidizertank vibration .................. 6-1hPropellant tank vibration qualification accelerations(a) Random, tank full ............... 6-15

Z.IAL

5.20-9

5.20-10

5.20-11

5.20-12

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xxviii

Figure

6.1-9

6. i-i0

6.1-11

9.3-1

9.3-29.3-39.3-49.3-59.3-69.3-?9.3-89.3-99.3-109.3-119.3-12

13.0-113.1-113.1-213.1-3

Page(b) Sine, tank full ................ 6-16(c) Sine, tank empty ............... 6-17Comparison of LTA-10R fuel tank acceleration spectral

density to qualification criteria ........ 6-18Comparison of LTA-10R descent engine acceleration

spectral density to qualification criteria .... 6-19Comparison of LTA-10R to LTA-3 sound pressure levels

spectra at lift-off ............... 6-20Apollo h launch abort areas and recovery force

structure .................... 9-27Booster and camera capsule recovery ........ 9-28

Mid-Pacific recovery zone and force deployment . . 9-29Primary recovery area and force deployment ..... 9-30Spacecraft and parachute .............. 9-31Forward heat shield ................ 9-32Spacecraft in flotation collar ........... 9-33Spacecraft after recovery ............. 9-34Left rendezvous window (inside view) ........ 9-35Right side window (inside view) .......... 9-36Camera capsules ................. 9-37S-IC/S-II Interstage ullage rocket motor

fairings ..................... 9-38Apollo space vehicle ................ 13-2Spacecraft 017 ................... 13-20Command module window assembly ........... 13-21Fuel cell power plant flow diagram ....... 13-22

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Figure13.1-413.1-513.1-6

13.1-7

C IA[Command and service module cryogenic schematic . . .Electrical power distribution subsystem ......Sequential events control subsystem functional

block diagram .................. 13-25Debris catcher configuration on spacecraft lunar

module adapter .................. 13-26Communications subsystem .............. 13-27Guidance and navigation subsystem functionalschematic _ o _

13.1-10 Stabilization and control subsystem functionalschematic .................... 13-2913.1-i1 Service module reaction control subsystem component

location .................... 13-3013.1-12 Command module reaction control subsystem ..... 13-3113.1-13 Service module reaction control subsystem propellant

feed subsystem .................. 13-3213.l-l& Command module reaction control subsystem propellant

distribution ................... 13-33

13.1-15 Service propulsion subsystem functional flowdiagram ..................... 13-3413.1-16 Environmental control subsystem schematic ..... 13-3513.1-17 Glycol coolant circuit ............... 13-3613.1-18 Vapor sensitive tape locations

(a) Tape locations 1-5 ..............(b) Tape locations 6-9 .............(c) Tape locations 10-15 .............

13.2-1 LTA-10R, descent stage ...............13.2-2 LTA-10R, ascent stage ...............

xxix

Page13-2313-24

13-3713-3813-39

13-4113-42

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XXX

Figure13.3-i

14. i-i

lh.i-2

14.2-i

14.2-2

TIAL

Saturn V launch vehicle with ApolloSpacecraft 017 ..................

Factory checkout flow for command and servicemodule at Downey .................

Prelaunch checkout flow for command and servicemodule at Kennedy Space Center ..........

Factory modification and checkout flow forLTA-10R at Bethpage ...............

Prelaunch checkout flow for LTA-10R for KennedySpace Center ...................

Page

13-46

14-2

14-3

14-5

14-5

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CO TIAL1.0 MISSION SUMMARY

1-1

The Apollo h mission was successfully accomplished on November 9,1967. This was the first Apollo mission utilizing a Saturn V launchvehicle with a lunar module test article (LTA-10R) and a Block I commandand service module (Spacecraft 017). The unmanned spacecraft waslaunched from complex 39A, Cape Kennedy, Florida. Lift-off occurred at1200:01 G.m.t., one second after the planned time, during the firstlaunch attempt. The spacecraft landed in the primary recovery area inthe Pacific Ocean near Hawaii approximately 8-1/2 hours later.

The principal objectives of the Apollo 4 mission were to demonstratethe structural and thermal integrity of the space vehicle and to verifyadequacy of the Block II heat shield design for entry at lunar returnconditions. These objectives were satisfactorily accomplished.

Performance of the spacecraft was satisfactory in all respects. Thelaunch phase of the flight was normal, with all planned events occurringwithin allowable limits. Strain gauge data indicated that no structuralfailures occurred and that structural loading was well within the capa-bility of the vehicle. Vibration data measured in the command moduleindicated that qualification vibration levels were not exceeded. Suffi-cient valid data were obtained on the spacecraft structure to enabledetermination of the thermal response during the launch phase and toverify the adequacy of the thermal analysis prediction techniques.

Performance of the emergency detection subsystem, operating in anopen-loop mode, was satisfactory. No conditions approaching manual orautomatic abort levels were encountered at any time during the launchphase.

The S-IVB stage inserted the spacecraft into an earth parkingorbit after approximately ll minutes of powered flight. After two rev-olutions in an earth parking orbit, the S-IVB stage was reignited fora simulated translunar injection burn. Shortly after this burn, thespacecraft separated from the S-IVB stage and the service propulsionsubsystem was ignited for a short-duration burn. No adverse effectswere noted as a result of starting the service propulsion subsystem, asplanned, in a zero-g environment with no reaction control subsystemullage maneuvers. This burn raised the apogee altitude to 9769 n_uticalmiles.

Following this service propulsion subsystem burn, the spacecraftwas aligned to a specific attitude to achieve a thermal gradient acrossthe command module heat shield. This spacecraft thermal orientationattitude, with the command module hatch window directly toward the sun

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._._ "_''_'TIAL

such that the conical surface of the crew compartment was perpendicularto the sun rays, was maintained for approximately _-1/2 hours. Theobjective of thermally conditioning the command module ablator prior toentry to induce circumferential thermal stresses and distortions on thecommand module was achieved. The cold-soak attitude was maintainedproperly throughout the coast phase although continuous venting from theenvironmental control subsystem water boiler produced a disturbancetorque which caused unsymmetrical limit cycles in the pitch and yaw axes.Block II thermal control coating degradation did occur, as evidenced bymeasured data exceeding the Block II nominal coating equilibrium tempera-tures. This degradation is attributed to materials used in the launchescape tower solid propellant jettison motor.

Following the cold-soak coast phase, the service propulsion sub-system was reignited for a long-duration burn to accelerate the space-craft to entry conditions that represent the most severe combination ofthe two extreme operational conditions that could possibly be achievedfrom a lunar return trajectory. Shortly after this burn, the commandmodule separated from the service module and the command module was ori-ented to entry attitude. Atmospheric entry, 400 000 feet, occurred atan inertial velocity of 36 629 ft/sec and a flight-path angle of minus6.93 degrees. The entry interface conditions were 210 ft/sec greaterand 0.20 degree shallower than had been predicted. The overspeed con-dition resulted from a longer than planned second service propulsionsubsystem burn. Because of the change in the entry conditions, the peakload factor of 7.27g was lower than the predicted 8.34g. These condi-tions did not affect the performance of the guidance system in achievingthe target.

The flight-derived total lift-to-drag ratio and total angle of attackwere well within the predicted uncertainty boundaries for the entirehypersonic flight regime. The flight derived data during the first entryare estimated to be a lift-to-drag ratio of 0.370 and a total angle ofattack of 15h.6 degrees.

Command module landing occurred within l0 nautical miles of theplanned landing point. The command module, the apex cover, and one ofthe three main parachutes were recovered by the prime recovery ship,U.S.S. Bennington. This was the first recovery of an Apollo parachute.

All spacecraft subsystems operated properly throughout the mission.There is no evidence of any functional anomalies that affected the mission.

The thermal protection subsystem survived the lunar entry environ-ment satisfactorily. Sufficient data were obtained to permit a thoroughevaluation of the thermal performance of the Block II thermal protection


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CO TIALsubsystem. Temperature data were within design limits. The maximumcalculated heating rate was 430 Btu/ft2/sec and the maximum calculatedheat load was 38 150 Btu/ft 2. These values are related to a referencepoint on the aft heat shield (location S/E = 0.9875, e = 90). Theexpected values were h22 Btu/ft2/sec and 34 750 Btu/ft 2, based on thelatest prediction method and the preflight trajectory. The expectedvalues before the mission were 594 Btu/ft2/sec and 37 777 Btu/ft 2. Thereason for the differences in the expected values is based on the use ofupdated radiative terms which are approximately 50 percent of thosepreviously used. Temperature data indicated that surface temperaturesapproached 5000 F. The maximum bondline temperature measured on theaft heat shield was 150 F. The surface erosion in the stagnation areaand other points on the aft heat shield was less than expected. Examina-tion of cores taken from the aft heat shield indicated a strong surfacechar and less-than-expected char penetration. The maximum char penetra-tion was 0.oo inch.

Performance of the guidance and control subsystems was equal to orbetter than preflight predictions. Analysis of navigation error propaga-tion during ascent, the service propulsion subsystem burns, and entryindicates that inertial measurement unit gyro drifts and accelerometerbiases and scale factors remained within specification tolerances. Allsequencing and computational operations performed by the Apollo guidancecomputer have been verified to have been correct.

Sequencing of the mission control programmer was satisfactorythroughout the mission.

Operation of the environmental control subsystem was satisfactory.The cabin pressure remained between 5.6 and 5.8 psia during the orbitalphase of the mission. The cabin temperature was maintained at a constant60 F during the mission, increasing to 68 F during entry.

Main dc power was satisfactorily supplied by three Block I fuelcells, augmented bythree auxiliary batteries during peak electricalloads. Extended zero-g operation of the fuel cells, inflight fuel cellreactant purge, and thermal control of the fuel cells were demonstrated.The cryogenic storage subsystem operation was also satisfactory. Extendedzero-g operation, equal depletion, pressure control, and stratificationcontrol were demonstrated.

The electrical power distribution subsystem operation was nominalthroughout the mission. All power switching occurred as planned andprogrammed.

All maneuvers using the reaction control subsystem thrusters werecompleted as planned. Satisfactory maneuver rates, accelerations, andtranslation velocity changes were attained.

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1-h CORALThe spacecraft and Manned Space Flight Network S-band comnunications

and spacecraft vhf communications were satisfactorily demonstrated. Gen-eral support from the NASA and Department of Defense network stations wasexcellent.

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\2.0

2-1

2.1 MISSION DESCRIPTION

The Apollo 4 space vehicle was launched from complex 39A at CapeKennedy, Florida, at 12:00:01 G.m.t. (07:00:01 a.m.e.s.t.) on Novem-ber 9, 1967. The launch azimuth was 72.0 degrees from true North. Thesequence of the major mission events is shown in table 2.1-I.

S-IC stage cutoff occurred nominally at 00:02:30.8 at an altitudeof approximately 34 nautical miles. S-II stage ignition occurred at anominal time of 00:02:32.2 with the stage burning for a period of 6 min-utes 7.6 seconds. The S-IVB stage was ignited and burned for 2 minutesearth parking orbit having an apogee of I01.i nautical miles and a peri-gee of 99.1 nautical miles.

After approximately two orbits (at 03:11:26.6), the S-IVB stage wasreignited to place the spacecraft into the simulated translunar trajec-tory. This burn was for a period of 4 minutes 59.7 seconds.

Approximately i0 minutes after the completion of the second S-IVBburn, a nominal spacecraft separation from the S-IVB was achieved. Oneminute 38.4 seconds after spacecraft separation, the first burn of theservice propulsion subsystem was initiated and, 16 seconds later, wascompleted satisfactorily. Upon completion of the burn, the spacecraftwas oriented to a cold-soak attitude which placed the thickest side ofthe command module heat shield away from the solar vector. During theapproximate 4.5-hour cold-soak period, the spacecraft coasted to itshighest apogee, 9769 nautical miles. Also, a 70-mm still camera wasphotographing the earth's surface once every 10.6 seconds during thisperiod. A total of 715 good-quality, high-resolution photographs weretaken during this period.

At 08:10:54.8, the service propulsion subsystem was ignited againto increase the spacecraft inertial velocity. The planned velocity was34 816 feet per second; however, the velocity achieved was 35 ll5 feetper second.

Two minutes 27.2 seconds after the completion of the second servicepropulsion subsystem burn, the command module was separated from theservice module, and the command module was oriented to the entry atti-tude.

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The entry phase of the flight was well within the conditions expect-ed. The lift-to-drag ratio obtained was 0.365 (0.015) compared withpreflight predictions of 0.350 (0.322 to 0.416). The spacecraft on itsmain parachutes was sighted from the recovery carrier, the U.S.S. Ben-nington, approximately 6 to 8 nautical miles from the recovery vessel.Landing occurred at 08:37:09.2 approximately lO nautical miles from theplanned landing point based on postflight reconstruction of the entrydata. Swimmers were deployed from helicopters and had a flotation col-lar secured around the spacecraft within 20 minutes. Recovery of thecommand module, apex heat shield, and one main parachute was effectedapproximately 2 hours 28 minutes after landing. This period was somewhatlonger than anticipated; however, sea conditions of 8-foot swells werethe cause of the longer recovery period.


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TIALTABLE 2.1-I.- APOLLO 4 MISSION EVENTS

2-3

Source

MSFC aMSFC aMSFC aMSFC aMSFC aCD0105MSFC aMSFC aMSFC aMSFC a

MSFC aTrajectory

dataMSFC aMSFC aCD0127CH4320CH4320Trajectory

dataCH4320 and

SP0661SP0661

CDO023Trajectory

dataCC0228CD0005CHOO2h

EventMission elapsed time,

hr:min:sec

Planned ActualLAUNCH PHASE

RANGE ZERO 12:00:01 G.m.t.

Lift-off (12:00:01.3 G.m.t.)Maximum dynamic pressureS-IC inboard engine cutoffS-IC outboard engine cutoffS-IC/S-II separation commandS-II engine ignition commandS-IC interstage jettisonLES jettisonS-II cutoffS-II/S-IVB separationS-IVB engine ignition co_and (first burn)S-IVB cutoff (first burn)

00:00:00.00:01:18.400:02:15.500:02:31.900:02:32.600:02:33.300:03:02.600:03:08.800:08:36.300:08:37.100:08:37.300:10:56.0

ORBITAL PHASE

Begin earth parking orbitBegin second earth orbit

S-IVB ignition (second burn)S-IVB cutoff (second burn)CSM/S-IVB separationSPS ignition (first burn)SPS cutoff (first burn)Apogee (planned 9890 n. mi.)

(actual 9769 n. mi.)SPS ignition (second burn)

SPS cutoff (second burn)

00:I1:06.001:38:20.0

03:11:33.503:16:39.903:26:42.803:28:20.i03:28:35.105:48:43.1

08:14:42.7

08:19:11.3ENTRY PHASE

cM/sM separation400 000 ft altitude

Drogue parachute deploymentMain parachute deploymentLanding

08:21:45.708:23:12.8

08:35:11.008:36:27.008:41:25.0

00:00:00.300:01:18.500:02:15.500:02:30.800:02:31.h00:02:32.200:03:01.400:03:07.200:08:39.800:08:40.500:08:40.700:11:05.6

00:ii:15.601:38:47

03:11:26.603:16:26.303:26:28.203:28:06.603:28:22.605:46:49.5

08:10:54.8

08:15:35.4

08:18:02.608:19:28.5

08:31:18.608:32:05.808:37:09.2

aMarshall Space Flight Center event times. All other times are from spacecraft measurementsor trajectory data.

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2-h

The four mission objectives relating to the spacecraft for theApollo 4 mission as assigned by the Office of Manned Space Flight arelisted below. All four have been satisfactorily accomplished.

1. Demonstrate the structural and thermal integrity and compati-bility of the launch vehicle and spacecraft. Confirm launch loads anddynamic characteristics.

2. Verify operation of the following subsystems: command moduleheat shield (adequacy of Block II design for entry at lunar return con-ditions), service propulsion subsystem (including no ullage start), andselected subsystems.

3. Evaluate the performance of the space vehicle emergency detec-tion subsystem (open-loop configuration).

4. Demonstrate mission support facilities and operations requiredfor launch, mission conduct and command module recovery.The detailed test objectives developed by MSC to support the four

mission objectives are listed in table 2.2-1 along with the degree ofaccomplishment and appropriate comments.

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LTABLE 2.2-I.- DETAILED TEST OBJECTIVES

2-5

No

1.1

1.2

1.4

ll.5

1.7

3.1

3.5

Category

N

M

M

M

M

Objective description

Demonstrate CSM/SLA/LTA/Saturn V struc-tural compatibility and determinespacecraft loads in a Saturn Vlaunch environment.

Determine the dynamic and thermal re-sponses of the SLA/CSM structure inthe Saturn V launch environment.

Determine the force inputs to the simu-lated LM from the SLA at the space-craft attachment structure in aSaturn V launch environment.

Obtain data on the acoustic and thermalenvironment of the SLA/simulated LMinterface during a Saturn V launch.

Determine vibration response of LM de-scent stage engine and propellanttanks in a Saturn V launch environ-ment.

Evaluate the thermal and structuralperformance of the Block II ThermalProtection System, including effectsof cold soak and maximum thermalgradient when subjected to the com-bination of a high heat load and ahigh heating rate representative oflunar return entry.

Demonstrate an SPS no ullage start.Determine performance of the SPS during

a long duration burn.Verify the performance of the SM-RCS

thermal control subsystem and enginethermal response in the deep spaceenvironment.

Degree ofaccomplishment

Satisfied.

Satisfied.

Partial.Only qualitativedata were obtained.

Satisfied.

Satisfied.

Satisfied.

Satisfied.Satisfied.

Satisfied.

M - denotes mandatoryP - denotes primary

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2-6 C IALTABLE 2.2-1.- DETAILED TEST OBJECTIVES - Continued

No.

3.6

3.8

3.9

5.2b

5.6

1

3.1c

3.1dPart

I3.1dPart

II

Degree ofCategory Objective description accomplishment

M Satisfied.

M

M

P

P

Verify the thermal design adequacy ofthe CM-RCS thrusters and extensionsduring simulated lunar return entry.

Evaluate the thermal performance of agap and seal configuration simulatingthe unified crew hatch design forheating conditions anticipated duringlunar return entry.

Verify operation of the heat rejectionsystem throughout the mission.

Evaluate the performance of the space-craft emergency detection subsystem(EDS) in the open-loop configuration.

Demonstrate the performance of CSM/MSFNS-band communications.

Measure the integrated skin and depthradiation dose within the commandmodule up to an altitude of at least2000 nautical miles.

Determine the radiation shieldingeffectiveness of the CM.

Demonstrate satisfactory operation ofCSM communication subsystem usingthe Block II type vhf omnidirectionalantennas.

Verify operation of the G&N systemafter subjection to the Saturn Vboost environment.

Verify operation of the EPS after beingsubjected to the Saturn V launch en-vironment.

Verify operation of PGS after beingsubjected to the Saturn V launchenvironment.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

M - denotes mandatoryP - denotes primaryS - denotes secondary


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CO NTIALTABLE 2.2-I.- DETAILED TEST OBJECTIVES - Concluded

2-7

No.

3.2a

3.2dPart

I3.2dPart

II3.3a

3.3c

3.3d

3.3e

Category

S

S

S

Objective description

Verify operation of the G&N in thespace environment after S-IVB separa-tion.

Verify operation of the EPS in thespace environment after S-IVB separa-tion.

Verify operation of the PGS in thetion.

Verify operation of the CM-RCS duringentry and recovery.

Verify operation of the G&N/SCS duringentry and recovery.

Verify operation of the EPS duringentry and recovery.

Verify operation of the ELS duringentry and recovery.

Obtain data via CSM-A/RIA communica-tions.

iGather data on the effects of a longduration SPS burn on spacecraftstability.

Obtain data on the temperature of thesimulated LM skin during launch.

Degree ofaccomplishment

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Satisfied.

Unknown. Evaluation ofsignal-to-noise ratiosand bit error rateanalyses have not yetbeen accomplished.

Satisfied.

Satisfied.

S - denotes secondary

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2-8 _NT_'A L

THIS PAGE INTENTIONALLY LEFT BLANK

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CON 'I:[AL3.0 TRAJECTORY DATA

3-1

This section contains a comparison of the planned and actual trajec-tories of the Apollo _ mission. The launch, orbital, and entry trajecto-ries referred to as planned in this section, are preflight calculatedtrajectories. The actual trajectories shown in this section are basedon the Manned Space Flight Network (MSFN) tracking data and the actualperformance and sequences as determined by airborne instrumentation.Marshall Space Flight Center (MSFC) supplied all trajectory data for thelaunch phase, parking orbit, and second S-IVB burn. An analysis of theApollo h trajectory through the second S-IVB burn may be obtained fromthe MSFC Apollo h Mission Report (ref. 1). The earth model for all tra-jectories and analyses of the MSFN trackers contained geodetic and grav-itational constants representing the Fischer ellipsoid. A ground trackorbit, and entry trajectories are presented in figures 3.1-2 to 3.1-5.

3.1 APOLLO h MISSION

The trajectory analysis presented herein is based on intermediatetrajectory data generated 21 days after the end of the mission. Thefinal trajectory data are being prepared at this time and will be re-leased in a supplemental report to this report on January 31, 1968.

3.1.i LaunchThe performance of all launch vehicle (AS 501) stages was reported

by MSFC as satisfactory. Mach 1 occurred 0.6 second earlier andO.1 n. mi. lower than planned. The maximum dynamic pressure was within3 psf of the expected, and 0.1 second later and 0.05 n. mi. lower thanplanned. A time history of the launch phase is presented in fig-ure 3.1-2; a comparison of the planned and actual trajectory parametersis contained in table 3.1-I.

3.1.2 Parking OrbitInsertion.- The first S-IVB cutoff occurred at 00:11:05.6, with

the time of insertion defined as 00:11:15.6. The insertion conditionspresented in table 3.1-II were obtained from the MSFC orbit determinationprogram which used first revolution tracking data from Bermuda, Carnarvon,and White Sands. Insertion elements are presented in table 3.1-III.

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3-2 C ALA time history of velocity, flight-path angle, and altitude is presentedin figure 3.1-3.

Venting trajectory.- The S-IVB venting polynomials from MSFC havebeen implemented into the MSC orbit determination program in order togenerate the best estimated trajectory during the parking orbit. Orbitalfits through the tracking data include revolution l, revolution 2, andrevolution 1 and 2. The state vectors which were solved for from thesefits agree within 500 feet in position and 1 ft/sec in velocity. Thebest trajectory, however, is a combination of fits through revolution1 and revolution 2, as shown by better agreement with vectors at inser-tion, lower tracker residuals, and small position and velocity differ-ences (177 feet and 0.5 ft/sec) at revolution interface.

Tables 3.1-IV and 3.1-V contain statistical summaries of the C-bandtracking data utilized in the fits. It should be emphasized that becauseseveral stations are being considered in each fit, the comparison ofactual and theoretical noise, and actual and theoretical bias limits area qualification of both the tracking data and data fit. An analysis ofall data has shown that the noise of the data was generally below thetheoretical value. The large values shown in the tables are the resultof the data fits and not noisy tracker data; however, the quality of thefit is a function of tracking data biases due to unsolved for parameters,which in these cases show that the fits are actually very good.

An evaluation of the real time computer complex (RTCC) navigationupdate vector prior to the second S-IVB ignition was performed in orderto explain the difference in the actual apogee achieved. The inflightRTCC navigation update vector, prior to the second S-IVB burn, was basedon a Bermuda vector at the start of revolution 2 (01:57:49). This vectorwas integrated to the update time (02:58:30.02) using a preflight ventingmodel which assumes a linear thrust. The postflight vector and the RTCCBermuda revolution 2 vector agree reasonably well (i.e., the differenceswere 1433 feet and 1 ft/sec. However, when the postflight vector wasintegrated to the update time, using the postflight venting polynomialsprovided by MSFC, the difference increased to 5.6 n. mi. in position andplus 51 ft/sec in velocity. Of the two vectors, the postflight vector_is more consistent with subsequent data, and indicates that the updatevector was high in velocity. This would make the Apollo guidance computerdetermine that it had achieved the proper apogee at cutoff (see sec-tion 3.1.3). It appears the large difference in the navigation updatevector is the result of the predicted venting model and the 1-hour propa-gation prior to the update.

In general, the unified S-band tracking data for the parking orbitwere not as good as expected. No known biases were found on the databy the RTCC, but the high-speed Doppler was considered too noisY to

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C iALbe useful. A detailed analysis of the unified S-band tracker data willbe contained in a supplement to this report.

Second S-IVB burn.- The S-IVB stage was ignited for the second timeat an elapsed time of 03:11:26.6 and burned for 299.7 seconds. Thetime of injection (10 seconds after cutoff) into the coast ellipse wasdefined as 03:16:36.3. The injection conditions presented in table 3.1-VIwere obtained by MSFC from the launch vehicle instrument unit data andfrom MSFN tracking data. These conditions agree with those obtained fromthe MSC preliminary BET within approximately 2 n. mi. in position and3 ft/sec velocity. Injection elements are presented in table 3.1-III.

3.1.3 Coast Ellipse----V ....... r _ ............. _=*1 l,..a--,..l,.,+a..a ,.a.,...i_c.L,a.c.a. La.,.jj. A ' ,. 2' .. ._ .. .U. .LJ ,. _" .. L G%_03:26:28.2, 602 seconds after the second S-IVB burn. Separation occurred

1.7 seconds after the RCS thrusters came on, followed by 8.h seconds of+X %ranslation. A comparison of planned and actual separation conditionsis presented in table 3.1-VI. The actual conditions were calculatedusing the Eastern Test Range (ETR) tracking data taken between the timeof the second S-IVB burn cutoff and first SPS burn ignition. These con-ditions were validated by integrating a vector before CSM/S-IVB separationthrough the normal separation sequence, and comparing the results at thetime of ignition for the first SPS burn with a vector obtained afterCSM/S-IVB separation. The two solutions agreed within 26h0 feet altitude,0.yh ft/sec velocity, and 0.1 degree flight-path angle, indicating thatactual separation was satisfactory. These conditions in table 3.1-VIagree with the MSFC separation conditions within approximately 1 n. mi.in position and 1 ft/sec in velocity.

Trackin6 analysis and reconstruction.- Tracking data for the coastellipse was satisfactory. Deviations between the C-band radar data andunified S-band orbit data were small. The best reconstructions haveshown position and velocity differences of only 3000 ft and 2 ft/sec,respectively, for the two solutions. Analysis shows that the correctionof several small data differences which have appeared during the analysisshould bring the two trajectories closer together. Table 3.1-VII pre-sents a statistical summary of the tracking residuals for both the C-bandradar and unified S-band radar data.

As noted above, several tracking data differences have been noted;primarily, a 300-foot jump in the C-band range residuals at 03:29:59 and05:12:29; and secondarily, a Carnarvon elevation bias on C-band radardata, and a time bias on Ascension unified S-band radar data. Theanalysis has not resolved these problems; however, this will not affectthe validity of the BET.

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3-4

SPS maneuvers and targeting.- The first and second SPS burn maneu-vers determined by the onboard guidance logic were nominal for the Apolloguidance computer state vectors indicated at the start of each burn andthe respective targeting parameters loaded in the computer erasablememory prior to launch. Each burn was reconstructed by using raw pulsedintegrating pendulous accelerometer data to determine the accelerationprofiles during the burns, and by using the engine and chamber pressuredata to simulate thrust buildup and tail-off. The SPS cases discussedin this analysis each include actual thrust characteristics. The firstand second SPS burns (SPS-1 and SPS-2, respectively) were simulated forthe navigational state vectors and the actual state vectors establishedfrom tracking data. The actual targeted values achieved for semilatusrectum (P) and orbit eccentricity (e) were determined from the trajecto-ries propagated from Eastern Test Range tracking vectors after SPS-1 andfrom the Guam tracking information after SPS-2. In both burns, it wasestablished that guidance cutoff was indicated for the times when theonboard targeting quantities were satisfied and when the simulated burntimes agreed very closely with Apollo guidance computer telemetry infor-mation.

Ignition for the first SPS burn occurred at 03:28:06.6 with guidancecutoff occurring 16.0 seconds later. The first SPS maneuver was a posi-grade guided burn targeted to an in-plane coast ellipse having an apogeeof 9899 n. mi., a P of 32 928 190 feet, and an e of 0.593874991. Thespacecraft pitch attitude at SPS-1 ignition was 72.2 degrees up from thelocal horizontal or 43.9 degrees up from the inertial velocity vector.The difference between actual and planned pitch attitude was about17 degrees which resulted from the second S-IVB burn providing a lowerearth-intersecting coast ellipse. The actual coast ellipse achieved bythe guided burn had an apogee of 9769 n. mi., a P of 32 833 060 feet, andan e of 0.59133947. Table 3.1-VI shows a state vector comparison of theplanned and actual conditions at SPS-1 ignition, guidance cutoff, andapogee. Figure 3.1-4 presents a time history of the SPS-1 burn for thetrajectory parameters inertial velocity, flight-path angle, and altitude.

To explain the differences between the planned and actual ellipsesachieved at the end of the first SPS burn, a number of SPS burn simula-tions were made. The burn characteristics and orbital parameters gener-ated for the various SPS first-burn simulations are presented intable 3.1-VIII.

Case I shows the actual conditions reconstructed from an ETR track-ing vector at the end of the first SPS burn. Case II is the best estimateof the first SPS burn which results from integrating an ETR trackingvector through a guidance and navigation subsystem guided burn targetedfor the actual P and e determined in Case I. In reconstructing the first3PS burn, the best estimate of the total thrust was calculated to be_l 106 pounds, from the known weight losses, burn duration, and total AV_chieved.

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TIAL 3-5Because the first SPS burn cutoff occurred when the Apollo guidance

computer calculated that the targeting conditions had been achieved, itwas apparent that the nominal P and e values loaded in the erasablememory had been satisfied. Cases III and IV show the results of a guid-ance cutoff on nominal targets for integrated ETR and AGC tracking vec-tors, respectively. As expected, Case IV indicated that the Apolloguidance computer burn logic yielded the desired P and e values and anapogee of 9890 n. mi. This apogee solution differs by 6.3 n. mi. fromthe Apollo guidance computer estimated solution. This difference isaccounted for by the fact that the apogee solution used in the Case IVsimulation of this burn took into account the effects of earth oblatenessand did not include any inertial measurement unit errors effective duringthe burn. Case III indicates that an additional 2.64 seconds of burntime would be required to achieve the planned P and e, if the onboardstate vector had agreed with the actual state vector at ignition of thefirst SPS burn. It follows that the difference between the value thatthe Apollo guidance computer computed for the state vector and the valueof the actual state vector yielded a burn which missed the planned apogeeby 121 n. mi. This difference in the actual and the Apollo guidancecomputer state vectors prior to the first SPS burn is attributed to thecombined effects of S-IVB venting, the navigation update, and inertialmeasurement unit errors.

Ignition for the second SPS burn occurred at 08:10:5h.8, with theground-commanded cutoff occurring 280.6 seconds later. The planned mis-sion called for the second SPS burn to accelerate the spacecraft to anentry interface velocity of 36 333 ft/sec and an entry interface flight-path angle of minus 7.13 degrees. The nominal target orbit parametersloaded in the Apollo guidance computer to accomplish the aforementionedobjectives included a semilatus rectum of 21 960 233 ft and an orbit ec-centricity of 0.9990992h. The actual target orbit achieved at the end ofthe second SPS burn had a semilatus rectum of h2 h88 012 ft and aneccentricity of 1.0221608. The spacecraft pitch attitude at ignitionfor the second SPS burn was 25.93 degrees below the inertial velocityvector or h9.13 degrees down from the local horizontal. These valuesagree within 0.7 degree of the planned pitch angles. Table 3.I-V con-tains a state-vector comparison of planned and actual conditions atsecond SPS burn ignition, guidance cutoff, and entry interface. Fig-ure 3.1-4 shows the planned and actual velocity, flight-path angle, andaltitude during the second SPS burn. It should he noted that the288-second deviation in planned and actual time of the ignition of thesecond SPS burn is the result of the lower apogee after the first SPSburn.

Four cases representing the simulated results of the second SPS burnare presented in table 3.1-VIII.

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3-6

Case V shows the actual conditions reconstructed from a Guam track-ing vector at the end of the second SPS burn. The second SPS burn enginecutoff command was accepted about 10.3 seconds after the guidance andnavigation engine cutoff command was indicated.

Case VI simulates the second SPS burn and resulted from integratinga tracking vector through a guided burn targeted to the actual P and edetermined in Case V. In reconstructing the burn, the best estimate ofthe actual thrust was calculated in the same manner as that for the firstSPS burn. Cases VII and VIII show the results of integrating a trackingvector and an Apollo guidance computer vector, respectively, through aguided burn targeted to the planned P and e. As can be seen from bothcases, if a guidance cutoff had been performed, the nominal target orbitand entry conditions would have been achieved.

The entry interface conditions for Case VIII indicated a velocityof 36 332.6 ft/sec and a flight-path angle of minus 7.13 degrees. Actualcutoff represented in Case V yielded a velocity of 36 544.6 ft/sec andflight-path angle of minus 6.93 degrees.

3.1.4 EntryThe planned and actual entry trajectories are shown in figure 3.1-5.As explained in section 3.1.3, the time shift between the actual and

planned curves can be attributed to the lower targeting achieved afterthe first SPS burn. The actual trajectory is based on a post-SPS-2 Guamradar vector, and the entry was generated by correcting the guidance andnavigation pulsed integrating pendulous accelerometer data for knowninertial measurement unit errors. Table 3.1-1X presents the planned andactual conditions at entry interface.

The entry interface conditions were 212 ft/sec greater and 0.20 de-gree shallower than planned. The off-nominal conditions did not affectthe performance of the guidance system in achieving the target.Table 3.1-X contains a comparison of the planned and actual values ofthe maximum entry parameters. Because of the change in the entry con-ditions, the peak load factor of 7.27g was slightly lower than predicted.

The analysis of the guidance and navigation subsystem has shown noanomalies and is discussed in section 5.15 of this report. The guidanceand navigation subsystem indicated a 2.2 n. mi. overshoot at drogueparachute deployment. The postflight reconstructed trajectory indicatesa 4.6 n. mi. undershoot at drogue parachute deployment.


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' ENTIA L 3-7TABLE 3.1-1.- LAUNCH PHASE PLANNED AND ACTUAL TRAJECTORY PARAMETERS

ICondition I Planned ActualS-IC Stage Inboard Engine Cutoff

Time from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg West ............Altitude, ft ................Altitude, n. mi ..............Space-fixed velocity, ft/sec .........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N . .

00:02:15.528.7580.08

159 00626

715122.9676.38

S-IC Stage Outboard Engine CutoffTime from range zero, hr :min:sec ......Geodetic latitude, deg North ...... Longitude, deg West ............Altitude, ft ...............Altitude, n. mi ..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N .

00:02:31.928.83

79.77208 691

348896

20.3375.62

00:02:15.528.7580.07

i62 86127

724123.2875.95

00:02:30.828.83

79.80208 990

348831

20.9675.29

,ENTIAL


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3-8 CTABLE 3.1-I .- LAUNCH PHASE PLANNED AND ACTUAL

TRAJECTORY PARAMETERS - Concluded

Condition PlannedS-II Stage Engine Cutoff

Actual

Time from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg West ............Altitude, ft ................

Altitude, n. mi ..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N . .

00:08:36.331.7065.72

622 510102

22 4820.52381.43

00:08:39.831.7265.67

631 050104

22 3560.64281.49

First S-IVB Stage Engine CutoffTime from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg West ............Altitude, ft ................Altitude, n. mi ..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N . .

00:10:56.032.6055.88

628 077103

25 561-0.00186.97

00:11:05.632.6455.h3

631 936104

25 5570.01587.21

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TABLE 3.1-11.- PARKING ORBIT PLANNED AND ACTUALTRAJECTORY PARAMETERS

3-9

Condition Planned ActualInsertion (S-IVB Stage Cutoff Plus i0 Seconds)

Time from range zero, hr:min:sec .....Geodetic latitude, deg North .......Longitude, deg West ...........Altitude, ft ................Altitude, n. mi .............Space-fixed velocity, ft/sec .......Space-fixed flight-path angle, deg ....Space-fixed heading angle, deg E of N.

00:11:06.032.6455.12

628 117103

25 57O0.00187.42

00:11:15.632.6754.67

631 670104

25 5640.01487.65

Second S-IVB Stage IgnitionTime from range zero, hr:min:sec .....Geodetic latitude, deg North .......Longitude, deg West ...........Altitude, ft ...............Altitude, n. mi .............Space-fixed velocity, ft/sec .......Space-fixed flight-path angle, deg ....Space-fixed heading angle, deg E of N.

03:11:33.531.9082.07

671 568ii0

25 551-0.00897.70

03:11:26.631.9582.33

668 045Ii0

25 547-0.00197.54


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3-10 C IALTABLE 3.l-II.- PARKING ORBIT PLANNED AND ACTUAL

TRAJECTORY PARAMETERS - Concluded

Condition Planned ActualSecond S-IVB Stage Cutoff

Time from range zero, hr:min:sec .....Geodetic latitude, deg North .......Longitude, deg West ............Altitude, ft ...............Altitude, n. mi .............Space-fixed velocity, ft/sec .......Space-fixed flight-path angle, deg .Space-fixed heading angle, deg E of N. .

03:16:39.927.9358.65

1 844 882304

3O 84O15.03

102.64

03:16:26.328.0359.36

1 766 542291

3g 88214.77

102.38


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T1A[TABLE 3.1-111.- ORBITAL ELEMENTS

3-11

Insertion(MSFC data)

injection

Coastellipse

ConditionApogee, n. mi ......Perigee, n. mi .....Period, min .......Inclination, deg ....Apogee, n. mi ......Perigee, n. mi .....Period, min .......Inclination, deg ....Apogee, n. mi ......Perigee, n. mi.Period, min .......Inclination, deg ....

Plannedi01.4lO0.O88.2

32.569410-45

306.230.319890-41

320.230.31

ActualI01.i99.188.3

32.579292-44

303.130.319769-45

316.630.31

IIA L


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3-12

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57/599

C 'I'IAL 3-13

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58/599

3-14 CO LTABLE 3.1-VI.- COAST ELLIPSE PLANNED AND ACTUAL

TRAJECTORY PARAMETERS

ICondition Planned J Actual

Injection (Second S-IVB Stage Cutoff Plus i0 Seconds)Time from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg West ............Altitude, ft ..............Altitude, n. mi ..............Spaced-fixed velocity, ft/sec .......Spaced-fixed flight-path angle, deg ....Space-fixed heading angle, deg E of N . .

CSM/S-IVB Separation

03:16:49.927.7758.08

1 925 302317

30 78615.29

103.02

03:16:36.327.8758.58

1 845 719304

30 82315.03

102.76

Time from range zero, hr:min:sec ......

Geodetic latitude, deg North ........Longitude, deg West ............Altitude, ft ................Altitude, n. mi ..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N . .

03:26:42.8

15.3524.94

8 082 3161330

26 18526.74

116.52

03:26:28.2

15.4325.10

7 948 4241308

26 23326.53

116.46


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TABLE3.1-VI.- COASTELLIPSEPLANNEDNDACTUALTRAJECTORYARAMETERSContinued

3-15

Condition PlannedFirst SPSIgnition

Time from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg West ............Altitude, ft ................Altitude, n. mi..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N . .

03:28:20.113.3621.33

9 248 1991522

25 45927.99

117.51First SPSCutoff

Time from range zero, hr :min:sec ......Geodetic latitude, deg North . . . .

Longitude, deg West ............Altitude, ft ..............Altitude, n. mi .............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N . .

03:28:35.1

13.0620.82

9 429 4961552

25 50728.44

117.64

Actual

03:28:06.613.4421.46

9 Ii0 5991499

25 5O427.80

117.46

03:28:22.6

13.1520.91

9 301 4021531

25 54728.30

117.59


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3-16 [TABLE 3.1-VI.- COAST ELLIPSE PLANNED AND ACTUAL

TRAJECTORY PARAMETERS - Continued

Condition Planned ActualApogee

Time from range zero, hr:min:see .. .....Geodetic latitude, deg South ........Longitude, deg East ............Altitude, ft ................Altitude, n. mi ..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N .

05:48:43.128.6936.39

60 092 3489890845O0.0

i00.38Second SPS Ignition

Time from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg East ............Altitude, ft ................Altitude, n. mi ..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....

Space-fixed heading angle, deg E of N . .

08:14:42.73.67

116.925 303 046

87328 235-23.14

59.87

05:46:49.528.6836.87

59 358 268976984O50.0

i00.38

08:10:54.83.45

117.495 340 719

87928 173-23.22

59.85


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TABLE3.I-VI.- COASTELLIPSEPLANNEDNDACTUALTRAJECTORYARAMETERSConcluded

3-17

Condition Planned ActualSecondSPSCutoff

Time from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg East ...........Altitude, ft ................Altitude, n. mi..............Space-fixed velocity, ft/sec .......Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N

08:19:11.312.64

131.932 279 663

37534 816-17.9862.16

CM/SMSeparationTime from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg East ............Altitude, ft ................Altitude, n. mi..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N . .

08:21:45.718.64

143.84900 749

14835 912-11.25

65.55

08:15:35.412.83

133.252 187 530

36O35 115-17.6462.21

08:18:02.618.62

144.74886 467

14636 138-11.07

65.49

CO .I TIAL


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3-18

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63/599

L 3-19

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64/599

3-20 CONFID LTABLE 3.1-IX.- ENTRY INTERFACE PLANNED AND ACTUAL

TRAJECTORY PARAMETERS

Condition Planned ActualTime from range zero, hr:min:sec ......Geodetic latitude, deg North ........Longitude, deg East ............Altitude, ft ................Altitude, n. mi ..............Space-fixed velocity, ft/sec ........Space-fixed flight-path angle, deg .....Space-fixed heading angle, deg E of N

08:23:12.821.90

151.58400 ooo

6636 333-7.1368.35

08:19:28.521.86

152.424OO 000

6636 545-6.9368.26


65/599

C TIAL 3-21TABLE 3. l-X.- SUMMARY OF PL_nNNED AND ACTUAL ENTRY PARAMETERS

Entry Conditions

ParameterMaximum entry velocity, ft/sec .......Maximum entry deceleration, g .......

Planned

Drogue Parachute Deployment Coordinates

Time from range zero, hr:min:sec . .....Latitude, deg North ...........Longitude, deg West ............

08:35:11.

apollo 4 mission report

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